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Development and Characterization of a Propulsion System for Cubesats Based on Vacuum Arc Thrusters

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Development and Characterization of a Propulsion System for Cubesats Based on Vacuum Arc Thrusters UNIVERSITÄT DER BUNDESWEHR MÜNCHEN Fakultät für Elektrotechnik und Informationstechnik Development and Characterization of a Propulsion System for CubeSats based on Vacuum Arc Thrusters Mathias Pietzka Vollständiger Abdruck der von der Fakultät für Elektrotechnik und Informationstechnik der Universität der Bundeswehr München zur Erlangung des akademischen Grades eines Doktor-Ingenieurs (Dr.-Ing.) genehmigten Dissertation. Vorsitzender: Prof. Dr.-Ing. Andreas Knopp Universität der Bundeswehr München 1. Gutachter: Prof. Dr.-Ing. Jochen Schein Universität der Bundeswehr München 2. Gutachter: Prof. Dr. Michael Keidar The George Washington University Die Dissertation wurde am 22.02.2016 bei der Universität der Bundeswehr München eingereicht und durch die Fakultät für Elektrotechnik und Informationstechnik am 01.07.2016 angenommen. Die mündliche Prüfung fand am 04.07.2016 statt. Abstract Within this PhD thesis the development of an innovative propulsion system for small satellites, especially so called CubeSats is described. The intended formation flight of such satellites requires a precise attitude control and fine positioning system. The mass (≤ 1.3 kg), size (100 × 100 × 100 mm3) and power (≈ 2 W) limitations of these satellites are challenging for an efficient propulsion system. One of the most promising concepts is the so called Vacuum Arc Thruster (VAT). Thereby, an electric discharge is initiated between two galvanically separated electrodes, an anode and a cathode, which leads to the erosion of the cathode surface. From the ejection of this material a thrust in the range of µN is produced at a specific Impulse of around 1000 s. However, up to now these values have been validated only under laboratory conditions without consideration of spacecraft limitations and requirements. Apart from the technical challenges of CubeSat integration, several principle issues have yet to be solved. These are the improvement of the unreliable ignition process, the feeding of the cath- ode material which is used as propellant, a uniform erosion of the cathode and the reduction of macroparticles which occur due to high localized heat loads. For the investigation of these effects and possible solutions, a comprehensive test environment was built consisting of a torsional thrust balance, an ion current density detector array and a high speed imaging system. Thereby it is possible to measure thrust, ion current density distribution and ion velocity as well as to observe the evolution of the plasma plume and the cathode spot movement optically in the range of µs. For arc initiation the so called “triggerless” ignition was implemented. Thereby, a conductive layer is applied to the insulation between anode and cathode which allows to reduce the necessary ignition voltage. However, this layer is eroded during thruster operation. An inwards sloped conical anode geometry enables a re-deposition of this layer with macroparticles from the cathode material and reduces spacecraft contamination. Titanium is chosen as cathode material because of its high feasibility as VAT propellant. A thrust-to-power ratio of around 12 µN/W given by literature was validated in laboratory tests. To facilitate a long lifetime, i. e. around 106 pulses for a typical ∆v of 7.5 m/s, a reliable feeding system was developed. Depending on the space, mass and power budget of the given satellite two feeding mechanisms can be used: A highly precise piezoelectric actuator which allows movements in the range of µm and a shape memory effect spring which uses the discharge current as energy source. The thus developed thruster has an outer diameter of 8.5 mm and a mass below 20 g. Due to the feeding system, a thruster length of around 30 mm is required which easily fits into the rail structure of a CubeSat. To meet the limitations and requirements of a typical CubeSat, a novel Power Processing Unit (PPU) was developed. Based on earlier attempts with an inductive energy storage this PPU uses an inductor to produce a voltage peak in the range of 1 kV for ignition. A capacitor is used as energy storage which allows pulse energies of up to 0.7 J and pulse lengths of up to 4 ms under the restrictions of a CubeSat. However, the maximal possible pulse repetition rate is limited to 1.4 Hz. To allow short periods of higher pulse energies (up to 2.7 J) a Li-Po battery was implemented. The ignition coil, which has been optimized in terms of a reduced production of parasitic magnetic fields, is bypassed during discharge to achieve a smooth discharge of the capacitor. An additional switch separates the VAT from the PPU to allow a charging of the capacitor without losses due to the conductive layer. Overall, the electric system has been miniaturized down to one PCB of 90×90 mm and a mass of 150 g. EMI is reduced by constructive measures like a splitting of power lines to several layers and the use of shielded wires. Four thrusters can be alternately operated by the PPU. For an improvement of thrust, specific impulse and erosion behavior a magnetic focusing system was implemented. Initially, the inductor of the PPU was used as field coil based on earlier work in literature. Due to the novel PPU design an additional coil with a low inductance but a high field strength was integrated into the circuit. Thereby, inductive losses have been reduced as well as the size and mass of the coil. It was found that, depending on the position of the coil, thrust can be increased by around 50 % or decreased by around 30 % and that the cathode spot movement is massively increased. Thereby, the local heat load and thus the production of macroparticles is reduced. This may help to prevent fusing of anode and cathode and to condition the re-deposition of the conductive layer. Zusammenfassung In der vorliegenden Arbeit wird die Entwicklung eines innovativen Antriebssystems für Kleinst- satelliten beschrieben. Der Formationsflug sogenannter CubeSats erfordert ein präzises Lage- und Positionsregelungssystem, das die Beschränkungen hinsichtlich Masse (< 1,3 kg), Größe (100 × 100 × 100 mm3) und Leistung (≈ 2 W) erfüllt. Eines der erfolgversprechendsten Konzepte ist der sogenannte Vacuum Arc Thruster (VAT). Dabei wird eine elektrische Entladung zwischen zwei galvanisch getrennten Elektroden erzeugt. Dies führt zur Erosion des Kathodenmaterials wodurch ein Schub im µm-Bereich bei einem spezifischen Impuls von ca. 1000 s erzeugt wird. Bei den meisten bisherigen Entwicklungen wurden die Beschränkungen und Anforderungen eines Satelliten nicht berücksichtigt. Abgesehen von den technischen Herausforderungen bei der Integration in einen CubeSat müssen einige grundlegende Probleme gelöst werden. Neben der Verbesserung des Zündprozesses und der Nachführung des als Reaktionsmasse genutzten Kathodenmaterials sind dies eine gleichmäßigere Erosion der Kathode und die Verringerung von Makropartikeln, die durch lokale Erhitzung entste- hen. Dazu wurde umfangreiche Testumgebung bestehend aus einer Schubmesswaage, einem Io- nenstromdichtedetektor und einem Hochgeschwindigkeitsbildaufnahmesystem aufgebaut. Dadurch ist es möglich Schub, Ionenstromdichteverteilung und Ionengeschwindigkeit zu messen bzw. die zeitliche Entwicklung der Plasmawolke und die Bewegung der Kathodenfußpunkte optisch im Be- reich von µs zu beobachten. Durch die sogenannte “triggerless” Zündung wird die erforderliche Zündspannung verringert indem eine leitfähige Schicht auf die Isolation zwischen Anode und Kathode aufgebracht. Allerdings wird diese Schicht beim VAT Betrieb erodiert. Eine nach innen konisch geformte Anode begünstigt nicht nur die Erneuerung dieser Schicht, sondern verringert auch die Kontamination des Satelliten. Als Kathodenmaterial wird Ti mit einem Schub-zu-Leistungsverhältnis von ca. 12 µN/W verwendet. Um eine lange Lebensdauer (ca. 106 Pulse für ein typisches ∆v von 7,5 m/s) zu erreichen, wurden zwei Nachführsysteme unter der Berücksichtigung der Raum-, Masse- und Leistungsbeschränkun- gen des Satelliten entwickelt: Zum einen ein hochpräziser Piezo-Aktuator, der Bewegungen im µm-Bereich erlaubt, und eine Formgedächtnisfeder, die mithilfe des Entladungsstromes aktiviert wird. Der hier entwickelte VAT hat einen Außendurchmesser von 8,5 mm, eine Länge von ca. 30 mm und eine Masse von 20 g. Des Weiteren wurde eine innovative Power Processing Unit (PPU) entwickelt. Die für die Zündung erforderliche Spannungsspitze im kV-Bereich wird dabei mit einer Induktivität erzeugt, während die Pulsenergie in einem Kondensator gespeichert wird. Unter den Leistungsbeschränkungen eines CubeSats sind damit Pulsenergien bis 0,7 J und Pulslängen bis 4 ms bei einer maximalen Pulsfre- quenz von 1,4 Hz möglich. Für eine kurzzeitige Erhöhung der Pulsenergie (bis 2,7 J) wurde ein Li-Po-Akku integriert. Die EMV-optimierte Zündspule wird während der Entladung überbrückt, um eine gleichmäßige Entladung des Kondensators zu erreichen. PPU und VAT werden während der Ladezeit voneinander getrennt um Verluste durch Stromfluss über die leitende Schicht zu ver- meiden. Die PPU wurde auf einer quadratischen Platine mit 90 mm Seitenlänge und einer Masse von 150 g untergebracht. Elektromagnetische Interferenzen werden durch konstruktive Maßnah- men reduziert. Die PPU ermöglicht den abwechselnden Betrieb von vier Triebwerken. Für die Verbesserung von Schub, spezifischen Impuls und Erosionsverhalten wurde ein magne- tisches Fokussiersystem integriert. Basierend auf früheren Veröffentlichungen wurde zunächst die PPU-Induktivität als Feldspule verwendet, die jedoch im weiteren Verlauf durch eine zusätzliche
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